U.S. patent application number 14/955261 was filed with the patent office on 2017-06-01 for thermal management of cmc articles having film holes.
The applicant listed for this patent is General Electric Company. Invention is credited to Ronald Scott Bunker, Kevin Robert Feldmann, Robert Charles Groves, II.
Application Number | 20170152749 14/955261 |
Document ID | / |
Family ID | 57421724 |
Filed Date | 2017-06-01 |
United States Patent
Application |
20170152749 |
Kind Code |
A1 |
Bunker; Ronald Scott ; et
al. |
June 1, 2017 |
Thermal Management of CMC Articles Having Film Holes
Abstract
Engine components are provided for a gas turbine engines that
generate a hot combustion gas flow. The engine component can
include a substrate constructed from a CMC material and having a
hot surface facing the hot combustion gas flow and a cooling
surface facing a cooling fluid flow. The substrate generally
defines a film hole extending through the substrate and having an
inlet provided on the cooling surface, an outlet provided on the
hot surface, and a passage connecting the inlet and the outlet. The
engine component can also include a coating on at least a portion
of the hot surface and on at least a portion of an inner surface
defined within the passage.
Inventors: |
Bunker; Ronald Scott; (West
Chester, OH) ; Feldmann; Kevin Robert; (Mason,
OH) ; Groves, II; Robert Charles; (West Chester,
OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
57421724 |
Appl. No.: |
14/955261 |
Filed: |
December 1, 2015 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2230/90 20130101;
F01D 9/065 20130101; F05D 2300/211 20130101; F23R 3/06 20130101;
F05D 2300/5024 20130101; F05D 2300/6033 20130101; F23R 2900/03042
20130101; F05D 2230/30 20130101; F05D 2300/15 20130101; C23C 28/321
20130101; F01D 5/186 20130101; Y02T 50/60 20130101; Y02T 50/6765
20180501; F05D 2260/202 20130101; C23C 28/345 20130101; F23R 3/007
20130101; Y02T 50/672 20130101; Y02T 50/676 20130101; F01D 5/288
20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F23R 3/00 20060101 F23R003/00; F23R 3/06 20060101
F23R003/06; F01D 9/06 20060101 F01D009/06 |
Claims
1. An engine component for a gas turbine engine generating hot
combustion gas flow, comprising: a substrate constructed from a CMC
material and having a hot surface facing the hot combustion gas
flow and a cooling surface facing a cooling fluid flow, wherein the
substrate defines a film hole extending through the substrate and
having an inlet provided on the cooling surface, an outlet provided
on the hot surface, and a passage connecting the inlet and the
outlet; and a coating on at least a portion of the hot surface and
on at least a portion of an inner surface defined within the
passage.
2. The engine component as in claim 1, wherein the CMC material has
a first thermal conductivity, and wherein the coating has a second
thermal conductivity, and further wherein the first thermal
conductivity is 10 times greater than the second thermal
conductivity or more.
3. The engine component as in claim 1, wherein the CMC material has
a first thermal conductivity, and wherein the coating has a second
thermal conductivity, and further wherein the first thermal
conductivity is 100 times greater than the second thermal
conductivity or more.
4. The engine component as in claim 1, wherein the coating is on at
least a portion of the hot surface, on at least a portion of an
inner surface defined within the passage, and on at least a portion
of the cold surface.
5. The engine component as in claim 2, wherein the coating has a
length extending away from respective film hole edge along the
cooling surface, the length being about 0.5 times to about 5 times
an inlet diameter in a direction of the hot combustion gas
flow.
6. The engine component as in claim 1, wherein the coating
perimetrically surrounds the hot surface around the outlet defined
in the hot surface.
7. The engine component as in claim 6, wherein the coating extends
across the entire hot surface of the substrate.
8. The engine component as in claim 1, wherein the coating has a
thickness defined from an external surface of the coating to the
hot surface of the substrate in a direction perpendicular to hot
combustion gas flow, the thickness being about 1000 .mu.m or
less.
9. The engine component as in claim 1, wherein the coating extends
around at least 50% of a downstream edge of the outlet.
10. The engine component as in claim 1, wherein the coating is over
at least a portion of the upstream inner surface and at least a
portion of the downstream inner surface.
11. The engine component as in claim 10, wherein the coating is on
the upstream inner surface through at least 50% of the depth from
the outlet to the inlet.
12. The engine component as in claim 10, wherein the coating is on
the upstream inner surface through at least 75% of the depth from
the outlet to the inlet.
13. The engine component as in claim 10, wherein the coating is on
the upstream inner surface from the outlet to the inlet.
14. The engine component as in claim 10, wherein the coating is on
the downstream inner surface through at least 50% of the depth from
the outlet to the inlet.
15. The engine component as in claim 10, wherein the coating is on
the downstream inner surface through at least 75% of the depth from
the outlet to the inlet.
16. The engine component as in claim 10, wherein the coating is on
the downstream inner surface from the outlet to the inlet.
17. The engine component as in claim 1, wherein the coating
completely covers all surfaces defined within the passage.
18. The engine component as in claim 1, wherein the coating has a
thermal conductivity that is less than the CMC substrate.
19. The engine component as in claim 1, wherein the coating is over
a thermal barrier coating forming the hot surface of the
substrate.
20. A gas turbine engine comprising: a compressor; a combustor; a
turbine; the engine component of claim 1.
Description
FIELD OF THE INVENTION
[0001] The present invention relates generally to ceramic matrix
turbine engine components, and more particularly, to a ceramic
matrix composite gas turbine engine component having small complex
features.
BACKGROUND OF THE INVENTION
[0002] In order to increase the efficiency and the performance of
gas turbine engines so as to provide increased thrust-to-weight
ratios, lower emissions and improved specific fuel consumption,
engine turbines are tasked to operate at higher temperatures. The
higher temperatures reach and surpass the limits of the material of
the components in the hot section of the engine and in particular
the turbine section of the engine. Since existing materials cannot
withstand the higher operating temperatures, new materials for use
in high temperature environments need to be developed.
[0003] As the engine operating temperatures have increased, new
methods of cooling the high temperature alloys comprising the
combustors and the turbine airfoils have been developed. For
example, ceramic thermal barrier coatings (TBCs) have been applied
to the surfaces of components in the stream of the hot effluent
gases of combustion to reduce the heat transfer rate, provide
thermal protection to the underlying metal and allow the component
to withstand higher temperatures. These improvements help to reduce
the peak temperatures and thermal gradients of the components.
Cooling holes have been also introduced to provide film cooling to
improve thermal capability or protection. Simultaneously, ceramic
matrix composites have been developed as substitutes for the high
temperature alloys. The ceramic matrix composites (CMCs) in many
cases provide an improved temperature and density advantage over
metals, making them the material of choice when higher operating
temperatures and/or reduced weight are desired.
[0004] CMCs have relatively low thermal conductivities and are thus
well suited for use in high temperature environments for long
periods of time. CMC components in the hot gas are heavily film
cooled, particularly in designs with otherwise uncooled trailing
edges. However, improved film cooling performance can decrease the
required cooling film flow and/or increase the durability of the
CMC component.
BRIEF DESCRIPTION OF THE INVENTION
[0005] Aspects and advantages of the invention will be set forth in
part in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0006] Engine components are generally provided for a gas turbine
engines that generate a hot combustion gas flow. In one embodiment,
the engine component includes a substrate constructed from a CMC
material and having a hot surface facing the hot combustion gas
flow and a cooling surface facing a cooling fluid flow. The
substrate generally defines a film hole extending through the
substrate and having an inlet provided on the cooling surface, an
outlet provided on the hot surface, and a passage connecting the
inlet and the outlet. The engine component also includes a coating
on at least a portion of the hot surface and on at least a portion
of an inner surface defined within the passage.
[0007] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following description and appended claims. The accompanying
drawings, which are incorporated in and constitute a part of this
specification, illustrate embodiments of the invention and,
together with the description, serve to explain the principles of
the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] A full and enabling disclosure of the present invention,
including the best mode thereof, directed to one of ordinary skill
in the art, is set forth in the specification, which makes
reference to the appended Figs., in which:
[0009] FIG. 1 shows a cross-sectional view of one embodiment of a
gas turbine engine that may be utilized within an aircraft in
accordance with aspects of the present subject matter;
[0010] FIG. 2 shows a side section view of an exemplary combustor
and a high pressure turbine of the engine from FIG. 1;
[0011] FIG. 3 is a schematic, sectional view through a film hole of
an exemplary engine component of the engine from FIG. 1 according
to one embodiment;
[0012] FIG. 4 is a schematic, sectional view through a film hole of
another exemplary engine component of the engine from FIG. 1
according to one embodiment;
[0013] FIG. 5 is top view of a plurality of film holes on an
exemplary engine compound according to one embodiment; and
[0014] FIG. 6 is top view of a plurality of film holes on another
exemplary engine compound according to one embodiment.
[0015] Repeat use of reference characters in the present
specification and drawings is intended to represent the same or
analogous features or elements of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0016] Reference now will be made in detail to embodiments of the
invention, one or more examples of which are illustrated in the
drawings. Each example is provided by way of explanation of the
invention, not limitation of the invention. In fact, it will be
apparent to those skilled in the art that various modifications and
variations can be made in the present invention without departing
from the scope or spirit of the invention. For instance, features
illustrated or described as part of one embodiment can be used with
another embodiment to yield a still further embodiment. Thus, it is
intended that the present invention covers such modifications and
variations as come within the scope of the appended claims and
their equivalents.
[0017] As used herein, the terms "first", "second", and "third" may
be used interchangeably to distinguish one component from another
and are not intended to signify location or importance of the
individual components.
[0018] The terms "upstream" and "downstream" refer to the relative
direction with respect to fluid flow in a fluid pathway. For
example, "upstream" refers to the direction from which the fluid
flows, and "downstream" refers to the direction to which the fluid
flows.
[0019] As used herein, the phrases "constructed of CMCs" and
"comprised of CMCs" shall mean components substantially constructed
of CMCs. More specifically, the CMC components shall include more
CMC material than just a layer or coating of CMC materials. For
example, the components constructed of CMCs may be comprised or
constructed substantially or entirely of CMC materials, including
greater than about 50, 60, 70, 80, 90, or 100 percent CMC
material.
[0020] Referring now to the drawings, FIG. 1 illustrates a
cross-sectional view of one embodiment of a gas turbine engine 10
that may be utilized within an aircraft in accordance with aspects
of the present subject matter, with the engine 10 being shown
having a longitudinal or axial centerline axis 12 extending
therethrough for reference purposes. In general, the engine 10 may
include a core gas turbine engine (indicated generally by reference
character 14) and a fan section 16 positioned upstream thereof. The
core engine 14 may generally include a substantially tubular outer
casing 18 that defines an annular inlet 20. In addition, the outer
casing 18 may further enclose and support a booster compressor 22
for increasing the pressure of the air that enters the core engine
14 to a first pressure level. A high pressure, multi-stage,
axial-flow compressor 24 may then receive the pressurized air from
the booster compressor 22 and further increase the pressure of such
air. The pressurized air exiting the high-pressure compressor 24
may then flow to a combustor 26 within which fuel is injected into
the flow of pressurized air, with the resulting mixture being
combusted within the combustor 26. The high energy combustion
products are directed from the combustor 26 along the hot gas path
of the engine 10 to a first (high pressure, HP) turbine 28 for
driving the high pressure compressor 24 via a first (high pressure,
HP) drive shaft 30, and then to a second (low pressure, LP) turbine
32 for driving the booster compressor 22 and fan section 16 via a
second (low pressure, LP) drive shaft 34 that is generally coaxial
with first drive shaft 30. After driving each of turbines 28 and
32, the combustion products may be expelled from the core engine 14
via an exhaust nozzle 36 to provide propulsive jet thrust.
[0021] It should be appreciated that each turbine 28, 30 may
generally include one or more turbine stages, with each stage
including a turbine nozzle and a downstream turbine rotor. As will
be described below, the turbine nozzle may include a plurality of
vanes disposed in an annular array about the centerline axis 12 of
the engine 10 for turning or otherwise directing the flow of
combustion products through the turbine stage towards a
corresponding annular array of rotor blades forming part of the
turbine rotor. As is generally understood, the rotor blades may be
coupled to a rotor disk of the turbine rotor, which is, in turn,
rotationally coupled to the turbine's drive shaft (e.g., drive
shaft 30 or 34).
[0022] Additionally, as shown in FIG. 1, the fan section 16 of the
engine 10 may generally include a rotatable, axial-flow fan rotor
38 that configured to be surrounded by an annular fan casing 40. In
particular embodiments, the (LP) drive shaft 34 may be connected
directly to the fan rotor 38 such as in a direct-drive
configuration. In alternative configurations, the (LP) drive shaft
34 may be connected to the fan rotor 38 via a speed reduction
device 37 such as a reduction gear gearbox in an indirect-drive or
geared-drive configuration. Such speed reduction devices may be
included between any suitable shafts/spools within engine 10 as
desired or required.
[0023] It should be appreciated by those of ordinary skill in the
art that the fan casing 40 may be configured to be supported
relative to the core engine 14 by a plurality of substantially
radially-extending, circumferentially-spaced outlet guide vanes 42.
As such, the fan casing 40 may enclose the fan rotor 38 and its
corresponding fan rotor blades 44. Moreover, a downstream section
46 of the fan casing 40 may extend over an outer portion of the
core engine 14 so as to define a secondary, or by-pass, airflow
conduit 48 that provides additional propulsive jet thrust.
[0024] During operation of the engine 10, it should be appreciated
that an initial air flow (indicated by arrow 50) may enter the
engine 10 through an associated inlet 52 of the fan casing 40. The
air flow 50 then passes through the fan blades 44 and splits into a
first compressed air flow (indicated by arrow 54) that moves
through conduit 48 and a second compressed air flow (indicated by
arrow 56) which enters the booster compressor 22. The pressure of
the second compressed air flow 56 is then increased and enters the
high pressure compressor 24 (as indicated by arrow 58). After
mixing with fuel and being combusted within the combustor 26, the
combustion products 60 exit the combustor 26 and flow through the
first turbine 28. Thereafter, the combustion products 60 flow
through the second turbine 32 and exit the exhaust nozzle 36 to
provide thrust for the engine 10.
[0025] FIG. 2 is a side section view of the combustor 26 and first
turbine 28 (i.e., the high pressure (HP) turbine) of the engine 10
from FIG. 1. The combustor 26 includes a deflector 76 and a
combustor liner 77. Adjacent to the turbine blade 68 of the turbine
28 in the axial direction are sets of axially-spaced, static
turbine vanes 72, with adjacent vanes 72 forming nozzles
therebetween. The nozzles turn combustion gas to better flow into
the rotating blades so that the maximum energy may be extracted by
the turbine 28. A cooling fluid flow C passes through the vanes 72
to cool the vanes 72 as hot combustion gas flow H passes along the
exterior of the vanes 72. A shroud assembly 78 is adjacent to the
rotating blade 68 to minimize flow loss in the turbine 28. Similar
shroud assemblies can also be associated with the LP turbine 32,
the LP compressor 22, or the HP compressor 24.
[0026] One or more of the engine components of the engine 10
includes a film-cooled substrate in which a film hole of an
embodiment disclosed further herein may be provided. Some
non-limiting examples of the engine component having a film-cooled
substrate can include the blades 68, vanes or nozzles 72, combustor
deflector 76, combustor liner 77, or shroud assembly 78, described
in FIGS. 1-2. Other non-limiting examples where film cooling is
used include turbine transition ducts and exhaust nozzles.
[0027] FIGS. 3 and 4 are a schematic, sectional view showing a
respective portion of exemplary engine components 80 formed from a
CMC substrate 82 defining at least one film hole 90 defined
therein. The engine component 80 may be an engine component of the
engine 10 from FIG. 1, and can be disposed in a flow of hot gas
represented by arrow H. A cooling fluid flow, represented by arrow
C may be supplied to cool the engine component. As discussed above
with respect to FIGS. 1-2, in the context of a turbine engine, the
cooling air can be first compressed air flow 54 which bypasses the
engine core 14, fluid from the LP compressor 22, or fluid from the
HP compressor 24.
[0028] The engine component 80 includes a substrate 82 having a hot
surface 84 facing the hot combustion gas flow H and a cooling
surface 86 facing the cooling fluid C. The substrate 82 may form a
wall of the engine component 80; the wall may be an exterior or
interior wall of the engine component 80. No matter the location or
type of component within the engine, the hot surface 84 of the
substrate 82 is exposed to hot gasses within the engine. The first
engine component 80 can define at least one interior cavity or
channel 88 comprising the cooling surface 86. The hot surface 84
may be an exterior surface of the engine component 80. In the case
of a gas turbine engine, the hot surface 84 may be exposed to gases
having temperatures in the range of 1000.degree. C. to 2000.degree.
C. Suitable materials for the substrate 82 include, but are not
limited to, steel, refractory metals such as titanium, or
superalloys based on nickel, cobalt, or iron, and ceramic matrix
composites. The superalloys can include those in equi-axed,
directionally solidified, and single crystal structures.
[0029] In one particular embodiment, the substrate 82 is
constructed from a ceramic matrix composite (CMC) material, which
is a non-metallic material having high temperature capability.
Exemplary CMC materials utilized for such substrate 82 may include
silicon carbide, silicon, silica or alumina matrix materials and
combinations thereof. Ceramic fibers may be embedded within the
matrix, such as oxidation stable reinforcing fibers including
monofilaments like sapphire and silicon carbide (e.g., Textron's
SCS-6), as well as rovings and yarn including silicon carbide
(e.g., Nippon Carbon's NICALON.RTM., Ube Industries' TYRANNO.RTM.,
and Dow Corning's SYLRAMIC.RTM.), alumina silicates (e.g., Nextel's
440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440
and SAFFIL.RTM.), and optionally ceramic particles (e.g., oxides of
Si, Al, Zr, Y and combinations thereof) and inorganic fillers
(e.g., pyrophyllite, wollastonite, mica, talc, kyanite and
montmorillonite). In one embodiment, the CMC material is formed
from a plurality of layers (e.g., about 4 to about 10 layers) in
its thickness.
[0030] The CMC material is formed into plies of material having
thicknesses of about 25 .mu.m to about 475 .mu.m, and most
typically about 125 to about 400 inches. The thicknesses of the
plies are usually dictated by the size of the fiber tows (filament
bundles) selected for use, and the thicknesses of the plies may
vary with fiber diameter. For most of the applications considered
herein, the plies are formed as two-dimensional woven fabric,
although one-dimensional fiber orientation may also be used.
However, the method of manufacturing the plies, laying up the plies
to form the component part and other parts manufacturing technology
used in the composite industry are not meant to limit the present
invention. The plies are laid up to form the shape of the article
being formed, the angles of the adjacent plies may vary depending
on the planar strength required. The components that can be made
using these CMC materials include but are not limited to turbine
blades, turbine vanes, turbine shrouds, and combustor liners,
casings, heat shields and diffusers. These hot section components
all benefit from the use of cooling air to provide sufficient
cooling to accomplish heat transfer during engine operation,
thereby extending their range of use.
[0031] The engine component 80 further includes one or more film
hole(s) 90 defined by the substrate 82 and extending through the
substrate 82 that provide fluid communication between the interior
cavity 88 and the hot surface 84 of the engine component 80. The
film hole 90 has an inlet 92 provided on the cooling surface 86 of
the substrate 82, an outlet 94 provided on the hot surface 84, and
a passage 96 connecting the inlet 92 and the outlet 94. The passage
96 is generally defined between the upstream inner surface 95 and
the downstream inner surface 97 within the passageway 96. While
only one film hole 90 is shown in FIG. 3, it is understood that the
engine component 80 may be provided with multiple film holes 90,
which be arranged in any desired configuration on the engine
component 80 (such as shown in FIGS. 5 and 6).
[0032] It is noted that, in any of the embodiments discussed
herein, although the substrate 82 is shown as being generally
planar, it is understood that that the substrate 82 may be curved
for many engine components 80. However, the curvature of the
substrate 82 may be slight in comparison to the size of the film
hole 90, and so for the purposes of discussion and illustration,
the substrate 82 is shown as planar. Whether the substrate 82 is
planar or curved local to the film hole 90, the hot and cooling
surfaces 84, 86 may be parallel to each other as shown herein, or
may lie in non-parallel planes.
[0033] During operation, the cooling fluid flow C is supplied to
the interior cavity 88, into the inlet 92, through the passageway
96 of the film hole 90, and out of the outlet 94 to create a thin
layer or film of cool fluid (for example, air drawn from the
compressor) over the hot surface 84, protecting it from the hot
combustion gas flow H. However, the film hole 90 has a relatively
high heat transfer coefficient on the upstream inner surface 95 and
the downstream inner surface 97, which can lead to the film hole 90
acting as an intense heat sink. However, since the CMC material has
a much lower thermal conductivity than metal (e.g., about half),
the film hole 90 and the flow within substantially increases
thermal gradients produced within the CMC locally, leading to
thermal stresses within the CMC substrate 82.
[0034] FIGS. 3 and 4 show a low conductivity coating 110 on at
least a portion of the upstream inner surface 95 and the downstream
inner surface 97. In certain embodiments, the low conductivity
coating 110 can be a thermal barrier coating (TBC) or an
environmental barrier coating (EBC). For example, the low
conductivity coating 110 can include a plurality of layers. For
example, the EBC can include a bond coat (e.g., comprising silicon
or silica), one or more layers formed from one or more rare earth
silicates (e.g., one or more of a mullite layer, a mullite-alkaline
earth aluminosilicate mixture layer, an yttrium monosilicate (YMS)
layer, an ytterbium doped yttrium disilicate (YbYDS) layer, a
barium strontium aluminosilicate (BSAS) layer, etc.), etc. The low
conductivity coating 110 is over any other coating (e.g., EBC
and/or TBC) that may be on the hot surface 84 of the substrate
82.
[0035] The low conductivity coating 110 can have a thermal
conductivity that is at least about 10 times less than the thermal
conductivity of the CMC substrate 100 so as to inhibit the
formation of thermal stresses within the substrate 100. That is,
the thermal conductivity of the CMC substrate is 10 times greater
(or more) than the thermal conductivity of the coating 110. For
example, the low conductivity coating 110 can have a thermal
conductivity that is at least about 50 times less than the thermal
conductivity of the CMC substrate 100. That is, the thermal
conductivity of the CMC substrate is 50 greater (or more) than the
thermal conductivity of the coating 110. As such, the coating 110
can protect the underlying CMC substrate 100, particularly in and
around the film hole, from forming a significant local temperature
gradient therein.
[0036] As shown in FIGS. 3 and 4, the low conductivity coating 110
is positioned on at least a portion of the hot surface 84 of the
substrate 82, on at least a portion of the inner surface 95, 97
defined within the passage (e.g., at least a portion of the
upstream inner surface 95 and/or at least a portion of the
downstream inner surface 97), and/or on at least a portion of the
cold surface 86 of the substrate 82. Although shown on a portion of
the hot surface 84 of the substrate 82, on the entire inner surface
95, 97 within the passage, and on a portion of the cold surface 86,
the low conductivity coating 110 can be applied as desired on each
of the respective surfaces 84, 86, 95, 97 (entirely or partially).
For example, the low conductivity coating 110 can extend on the
upstream inner surface 95 and/or the downstream inner surface 97
from the outlet 94 to the inlet 92. In one particular embodiment,
the coating 110 completely covers all surfaces defined within the
passage.
[0037] The low conductivity coating 110 extends, in one embodiment,
around at least 50% of an upstream edge 102 of the outlet 94 (e.g.,
around at least 75% of the upstream edge of the outlet), such as
shown in FIG. 6. Similarly, the low conductivity coating 110 can
extend, in one embodiment, around at least 50% of a downstream edge
104 of the outlet 94 (e.g., around at least 75% of the downstream
edge of the outlet). For example, the low conductivity coating 110
perimetrically surrounds the outlet 94 defined in the hot surface
84, as shown in FIG. 5. Similarly, the low conductivity coating 110
perimetrically surrounds around the inlet 96 defined in the cold
surface 86.
[0038] In certain embodiments, the coating 110 can be an extension
of any or all layers of a thermal barrier coating (TBC), an
environmental barrier coating, an adhesion compliance coating, etc.
that extends across the entire hot surface 84. However, in other
embodiments, the low conductivity coating 110 can extend only
partially from the film hole. For example, the low conductivity
coating 110 can have an average length extending away from
respective film hole edge (e.g., the upstream edge 102 and/or the
downstream edge 104) that is about 0.5 times to about 10 times
(e.g., about 0.5 times to about 5 times) the outlet diameter in a
direction of the hot combustion gas flow H measured from an
upstream inner surface 102 to a downstream inner surface 104. As
shown in the exemplary embodiments of FIGS. 3 and 4, the outer ends
113, 115 (away from the respective film hole edge) of the low
conductivity coating 110 are tapered to the hot surface 84 so as to
lessen any impact on the airflow across the film hole 90. However,
in other embodiments, the outer ends 113, 115 may have a different
orientation with respect to the outer surface 84, such as
perpendicular to the hot surface 84, curved, stepped, etc.
[0039] Similarly, the coating 110 can extend across the entire
cooling surface 86, or can extend only partially from the inlet 92
of the film hole 90. For example, the low conductivity coating 110
can have an average length extending away from respective film hole
edge of the inlet 92 (e.g., the upstream edge and/or the downstream
edge) that is about 0.5 times to about 10 times (e.g., about 0.5
times to about 5 times) the inlet diameter in a direction of the
hot combustion gas flow H measured from an upstream inner surface
to a downstream inner surface. As shown in the exemplary
embodiments of FIGS. 3 and 4, the outer ends can be tapered to the
cooling surface 86 or may have a different orientation with respect
to the cooling surface 86, such as perpendicular to the cooling
surface 86, curved, stepped, etc.
[0040] The low conductivity coating 110 has, in particular
embodiments, a thickness on the respective surface of the substrate
82 that is about 10 .mu.m to about 1500 .mu.m, but may vary in
thickness depending on the location of the coating. For example,
the low conductivity coating 110, in one embodiment, has a
thickness h on the hot surface 84 that is defined from an external
surface 106 of the low conductive coating 110 to the hot surface 84
of the substrate 82 in a direction perpendicular to hot combustion
gas flow H. In one embodiment, the thickness h is about 1500 .mu.m
or less, preferably about 25 .mu.m to about 500 .mu.m. The
thickness h', h'' on the inner surfaces 95, 97, respectively,
within the film hole 90 is, in most embodiments, about 10 .mu.m to
about 130 .mu.m. As such, in one particular embodiment, the
thickness h of the low conductivity coating 110 on the hot surface
84 is greater than the thickness h', h'' on the inner surfaces 95,
97.
[0041] The film holes 90 can have any shape as desired, such as a
tapered cross-section (e.g., expanding or contracting).
Additionally, the film holes 90 can include any features therein,
such as including a metering section (for metering of the mass flow
rate of the cooling fluid flow C) and a diffusing section (in which
the cooling fluid C may expand to form a wider cooling film),
and/or other features.
[0042] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they include structural elements that do not
differ from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
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